1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a rotor blade and stator vane arrangement to reduce pressure side vortices on the vanes.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes multiple stages of stator vanes and rotor blades in the turbine section that are exposed to a high temperature gas flow. The stator vanes guide the hot gas flow into the adjacent and downstream rotor blades in order to increase the performance of the turbine.
As the hot gas flow entering the turbine with a boundary layer thickness and reacts with the leading edge of the vane airfoil, a horseshoe vortex 10 separates into pressure side vortex 11 and a suction side vortex 12 as seen in FIG. 1. Initially, the pressure side vortex 11 sweeps downward and flows along the airfoil pressure side forward fillet region first. Then, due to the hot gas flow channel pressure gradient occurring between the pressure side to the suction side, the pressure side vortex 11 migrates across the hot flow passage and ends up at the suction side (13 in FIG. 2) of the adjacent airfoil. As the pressure side vortex 11 rolls across the hot gas flow channel, the size and strength of the passage vortex 15 becomes larger and stronger. Since the passage vortex 15 is much stronger than the suction side vortex 12, the suction side vortex 12 flows along the airfoil suction side fillet and acts as a counter vortex for the passage vortex 15. These vortices formation for a boundary layer entering the turbine airfoil can be seen in FIG. 1. As a result of the vortices flow phenomena, some of the hot core gas flow from the upper airfoil span is transferred toward close proximity to the endwall and thus creates a high heat transfer coefficient and high gas temperature region at the airfoil fillet region.
As shown in FIG. 1, the resulting forces drive the stagnated flow 21 that occurs along the airfoil leading edge towards the region of lower pressure at the intersection of the airfoil and endwall. This secondary flow flows around the airfoil leading edge fillet and endwall region. This secondary flow 21 then rolls away from the airfoil leading edge and flows upstream along the endwall against the hot core gas flow 20 as seen in FIGS. 2 and 3. As a result, the stagnated flow forces acting on the hot core gas and radial transfer of hot core gas flow from upper airfoil span toward close proximity to the endwall and creates a high heat transfer coefficient and high gas temperature region at the intersection location.
In the prior art, injection of film cooling air at discrete locations along the horseshoe vortex region is used to provide the cooling for this design. However, there are many drawbacks for this type of film blowing injection cooling and includes the following. A high film effectiveness level is difficult to establish and maintain in the high turbulent environment and high pressure variation region such as the horseshoe vortex region. Film cooling is very sensitive to pressure gradient. The mainstream pressure variation is very high at the horseshoe vortex location. The spacing between the discrete film cooling holes and areas immediately downstream of the spacing are exposed to less or no film cooling air. Thus; these areas are more susceptible to thermal degradation and metal over-temperature and shorten the life of the airfoil.
FIGS. 4 and 5 shows a stator vane with a leading edge and a pressure side of the airfoil, with an endwall having a fillet where the endwall merges into the airfoil. The horseshoe vortex region 23 is located in the leading edge and porous plug cooling holes 24 open onto the endwall in this location. Discrete pressure side film cooling holes 25 are on the airfoil just above the fillet. FIG. 4 shows the suction side film cooling holes 26 along the trailing edge section.
For the rotor blade trailing edge root section of the prior art, due to the hot gas migration from blade upper span down to the trailing edge versus the platform region, the blade aft fillet region experiences hotter gas temperature. Also, at the blade trailing edge fillet location, due to the stress concentration issue, the cooling slot for the airfoil trailing edge root section cannot be located low enough into the blade root section fillet region to provide proper convective cooling. Cooling of this particular blade trailing edge root fillet region becomes especially difficult.